The present invention relates generally to cooled turbine elements for gas turbine engines, and more particularly, to a method of lowering a stress in a cooled turbine element and the element made thereby.
FIG. 1 illustrates a portion of a gas turbine engine, generally designated by the reference number 10. The gas turbine engine 10 includes cooled turbine elements such as a high pressure turbine nozzle 12, a high pressure turbine blade (generally designated by 14), and a first stage low pressure turbine nozzle 16. As illustrated in FIG. 2, each of these cooled elements (e.g., blade 14) includes one or more airfoils 20, and one or more flowpath boundary members (e.g., a blade platform, generally designated by 22). In the case of the turbine blade 14, the element also includes a conventional dovetail 24 for connecting the blade to a turbine disk 26 (FIG. 1), and a shank 28 extending between the dovetail and the blade platform 22. Interior cooling passages 30 extend from openings (not shown) at the inner end of the blade dovetail 24 to cooling holes 32 in the airfoil 20. The passages 30 convey cooling air through the blade to remove heat from the blade. The cooling air passing through the cooling holes 32 in the airfoil 20 provides a film cooling barrier around the exterior surface of the airfoil.
Each flowpath boundary member 22 has a flowpath face 34 which faces the flowpath of the engine 10 and an outside face 36 opposite the flowpath face. As will be appreciated by those skilled in the art, the flowpath face 34 of each flowpath boundary member 22 runs hotter than the outside face 36 during engine operation. This difference in temperature results in the flowpath face 34 tending to grow more as a result of thermal growth than the outside face 36. Because the boundary member 22 is constrained by the airfoil 20, the tendency for the flowpath face 34 to grow more than the outside face 36 produces thermal stresses in the boundary member and the airfoil. More particularly, tensile stresses are produced in a trailing edge 38 of the airfoil 20 due to the tendency for the flowpath face 34 to grow more than the outside face 36. Experience has shown that fatigue cracks form and propagate as a result of the tensile stresses in the trailing edge 38 of the airfoil 20, resulting in a shortened life of the blade 14. Thus, there is a need for a method of lowering these stresses in colled turbine elements.
Briefly, apparatus of this invention is a cool turbine element for use in a flowpath of a gas turbine engine. The element comprises an airfoil having a pressure side and a suction side opposite the pressure side. The pressure side and the suction side extend axially between a leading edge and a trailing edge opposite the leading edge and radially between an inboard end and an outboard end opposite the inboard end. Further, the element comprises a flowpath boundary member extending laterally from at least one of the inboard end and the outboard end. The boundary member has a flowpath face and an outside face opposite the flowpath face. The outside face runs cooler than the flowpath face during engine operation thereby creating a tendency for the member to deflect in a direction away from the flowpath face and causing a thermally induced tensile radial stress in a region of the trailing edge of the airfoil. In addition, the element comprises an interior cooling passage extending through the airfoil from a cooling air source for transporting cooling air through the airfoil and at least one cooling hole extending from the interior cooling passage to an opening located on one of the suction side and the pressure side in an area upstream from the stressed region of the trailing edge to cool the area to a temperature below that of the trailing edge so that the airfoil thermally deflects during engine operation to a shape corresponding to that of the flowpath boundary member thereby lowering the thermally induced tensile radial stress in the airfoil at the trailing edge thereof.
In another aspect, the invention includes a method of lowering a tensile stress at a trailing edge of an airfoil of a cooled blade adjacent a platform of the blade. The method comprises the step of forming at least one cooling hole in the airfoil from an interior cooling air passage to an exterior surface of the airfoil to deliver cooling air to the exterior surface to cool an area of the exterior surface immediately adjacent the cooling hole thereby shifting tensile thermal loading from regions of the airfoil adjacent the area of the exterior surface to the cooled area.
In yet another aspect, the present invention includes a method of lowering a thermal stress at a trailing edge of an airfoil of a cooled turbine blade adjacent a platform of the blade. The method comprises the step of forming at least one cooling hole positioned upstream from the trailing edge of the airfoil and extending from an interior cooling air passage to an exterior surface of the airfoil for delivering cooling air to the exterior surface to cool the airfoil in an area of the exterior surface upstream from the trailing edge so that a thermal deflection of the airfoil more closely corresponds to a thermal deflection of the platform thereby lowering thermally induced stresses in the airfoil at the trailing edge thereof.
Other features of the present invention will be in part apparent and in part pointed out hereinafter.